Compressor stator with leading edge fillet

ABSTRACT

A compressor of a gas turbine engine includes a rotor and a stator located downstream of the rotor. The stator has a plurality of vanes each with an airfoil extending span-wise between a root proximate an inner hub of the stator and a tip. A fillet is disposed at the leading edge of the root of the airfoil, and extends between a pressure side surface of the airfoil and the inner hub.

TECHNICAL FIELD

The application relates generally to stators for gas turbine engines, and more particularly, to compressor stators.

BACKGROUND

The fans of many turbofan engines have fan blades with a large slope at the root of the fan airfoils and a large change in radius, from the leading edge to the trailing edge, at the fan blade roots. These properties may provide certain aerodynamic advantages. However, when fan chord is minimized for engine length/weight reasons, or fan blade root is thickened for structural reasons, the resulting high slope at the fan blade root can compromise the downstream fan root flow.

Consequently, the flow downstream of the fan blade root in such fans can carry large circumferential wake and thick end wall boundary layers. This can cause high incidence angle of the flow at the downstream stators, which may cause undesirable effects, such as the initiation of premature stall on the fan core stators downstream of the fan.

SUMMARY

There is accordingly provided a compressor for a gas turbine engine, the compressor comprising: a rotor and a stator located downstream of the rotor, the stator comprising a plurality of vanes, each of the vanes having an airfoil extending generally radially from a root proximate an inner hub of the stator to a radially outer tip, a span of the airfoil defined between the root and the tip, the airfoil having a leading edge, a trailing edge, and a chord extending between the leading edge and the trailing edge to define a chord length, the airfoil having a pressure side surface and a suction side surface each extending on opposite sides of the airfoil between the leading edge and the trailing edge, and a fillet disposed at the root of the airfoil, the fillet extending upstream from the leading edge of a remainder of the airfoil, the fillet extending away from the pressure side surface a greater distance than the fillet extends away from the suction side surface.

There is also provided a turbofan engine comprising a fan and a casing defining a bypass duct surrounding an engine core defining an annular gas passage, a fan stator disposed within the engine core downstream of the fan, the fan stator including a plurality of vanes circumferentially spaced-apart around a circumference of fan stator within the annular gas passage, each of the vanes having an airfoil extending between a root and a tip spaced apart by a span of the airfoil, the airfoil having a leading edge and a trailing edge spaced apart along a chord line by a chord length of the airfoil, a pressure side surface and a suction side surface respectively extending on opposite sides of the airfoil between the leading edge and the trailing edge, and a leading edge fillet disposed at the root of the airfoil on the pressure side surface.

There is further provided a gas turbine engine comprising: a compressor with a rotor and a stator located downstream of the rotor, the stator having a plurality of vanes each having an airfoil extending span-wise between a root proximate an inner hub of the stator and a tip, the airfoil extending chord-wise between a leading edge and a trailing edge, a fillet disposed at the leading edge of the root of the airfoil and extending between a pressure side surface of the airfoil and the radially inner hub.

BRIEF DESCRIPTION OF THE DRAWINGS

Reference is now made to the accompanying figures in which:

FIG. 1 is a schematic cross-sectional view of a gas turbine engine;

FIG. 2 is a schematic side elevational view of a vane of a compressor stator of the gas turbine engine of FIG. 1;

FIG. 3 is a schematic cross-sectional view of the vane of the compressor stator of FIG. 2, taken through line 3-3 in FIG. 2;

FIG. 4 is a schematic cross-sectional view of the vane of the compressor stator of FIG. 2, taken through line 4-4 in FIG. 2;

FIG. 5A is a partial side perspective view of the vane of the compressor stator of FIG. 2; and

FIG. 5B is a partial front perspective view of the vane of the compressor stator of FIG. 2.

DETAILED DESCRIPTION

FIG. 1 illustrates a gas turbine engine 10 of a type preferably provided for use in subsonic flight, generally comprising in serial flow communication a compressor 13 for pressurizing the air, a combustor 14 in which the compressed air is mixed with fuel and ignited for generating an annular stream of hot combustion gases, and a turbine section 15 for extracting energy from the combustion gases.

The compressor 13 includes one or more axial compressor stages 16 within the core 22 of the gas turbine engine 10. Although the gas turbine engine 10 may be of different types (turboprop, turboshaft, turbofan, etc.), in the embodiment depicted in FIG. 1 the gas turbine engine 10 is a turbofan engine, and therefore includes a fan 12, upstream of the engine core 22, through which ambient air enters the engine. The air propelled downstream from the fan 12 is split into a bypass duct 24 and the engine core 22.

Each compressor stage 16 within the engine core 22 includes one or more rows of compressor stators 17 located immediately downstream of a row of compressor rotors 18. A fan core stator 17A, located downstream of the bladed rotor of the fan 12, is also disposed within the engine core 22, near an inlet thereof and upstream from the axial compressor stages 16. It is understood that the fan 12 forms part of the compressor 13 of the engine 10, and comprises a first stage, low pressure compressor. Accordingly, the fan core stator 17A and the compressor stators 17 of the downstream axial compressor stages 16 all constitute compressor stators, as defined herein, and may therefore have the compressor stator vanes 20 as will be defined hereinbelow.

Each of the compressor stators 17,17A is a non-rotating component that guides the flow of pressurized air downstream from the compressor core rotors 18 and/or the fan 12. The compressor rotors 18 and the fan 12 rotate about a longitudinal center axis 19 of the gas turbine engine 10 to perform work on the air.

Each of the compressor stators 17,17A comprises a plurality of stator vanes 20. Each stator vane 20 is a stationary body that diffuses the airflow impinging thereon, thereby converting at least some of the kinetic energy of the incoming airflow into increased static pressure. The stator vanes 20 also redirect the airflow toward the next downstream compressor rotor 18 and/or to the combustor (in the case of the most-downstream compression stage 16).

With reference now to FIGS. 2-5B, the stator vanes 20 of the compressor stators 17,17A of the compressor 13 will now be described in further detail. It is understood that only one, or any two or more, of the stages of the compressor 13 may have the stator vanes 20 as described hereinbelow. For example, only the fan core stator 17A may be formed with the stator vanes 20 as described herein. Alternately, only one or more of the stator vanes 17 of the compressor 13 may have the stator vanes 20 as described herein. Alternately still, both the fan core stator 17A and the stator vanes 17 may all have the stator vanes 20 as described herein. The terms “upstream” and “downstream” as used herein are made with reference to the flow F of air through the compressor stators 17,17A and thus the airfoil F over each vane 20 thereof.

Referring first to FIGS. 2-3, each stator vane 20 (or simply “vane” 20) of the present disclosure includes an airfoil 21 shaped and sized to effect the above-describe functionality. The airfoil 21 extends generally radially between a root 26, disposed adjacent to a radially inner hub 27 of the compressor stator 17,17A, and a distal tip 28, disposed adjacent to an outer shroud 29 of the compressor stator 17,17A. The airfoils 21 define a span S of the vane 20 which extends between the roots 26 and the tips 28. The airfoil 21 therefore extends substantially in a radial direction (i.e. in a direction that generally extends parallel to a radial line from the center axis 19 of the gas turbine engine 10), which extends between the root 26 and the tip 28, and which defines the radial span S of the airfoil 21. This may also be referred to as the height of the vanes 20.

The chord length C of the airfoil 21 is defined between a leading edge 30 of the airfoil 21, and a trailing edge 32 of the airfoil 21. The extent of the airfoil 21 along its chord is therefore defined between the leading edge 30 and the trailing edge 32, and is referred to herein as the chord length C. In the depicted embodiment, the chord length C is the length of the chord line, which may be thought of as a straight line connecting the leading and trailing edges 30 and 32.

The airfoil 21 also includes a pressure side surface 34 and a suction side surface 36, disposed on opposite sides of the airfoil and each extending between the leading edge 30 and trailing edge 32.

Referring still to FIGS. 2-3, the vane 20 further includes a fillet 40 disposed at the root 26 of the airfoil 21, proximate the leading edge 30. Each of the airfoils 21 accordingly has a root 26 with an asymmetrical fillet 40 on a leading edge pressure side of the airfoil. The asymmetrical filet 40 is such that it is larger on the pressure side of the airfoil than on the suction side of the airfoil. The term “larger” as used herein in this context is understood to mean a filet that is any one or more of taller, longer, wider, greater radius, thicker, greater volume, greater mass, or the like, and/or any combination of these characteristics. In the embodiment of FIGS. 2-3, for example, the filet 40 on the pressure side 34 of the airfoil 21 protrudes outwardly from the pressure side surface of the airfoil a greater extent than does any filet (if one is present at all) located on the suction side 36 of the airfoil, and is therefore said to be larger than any filet on the suction side 36. (In this embodiment, however, the suction side 36 is substantially free of any filet.)

The fillet 40 may extend between the airfoil 21 and the radially inner vane platform 27, as seen in FIGS. 2 and 3. In the depicted embodiment, the fillet 40 is disposed exclusively on the pressure side 34 of the airfoil. While the fillet 40 may extend onto the suction side surface 36 of the airfoil 21 in certain alternate embodiments, in all cases the fillet 40 is larger on the pressure side surface 34 of the airfoil 21 than on the suction side surface 36, such a way that the fillet 40 is configured to provide a greater aerodynamic effect on the pressure side of the airfoil than on the suction side thereof. The term “larger” as used in this context is understood to mean any one of wider, higher, longer and/or thicker (or other corresponding sized-based physical properties). In the depicted embodiment, the fillet 40 is disposed only on the pressure side 34 only of the airfoil 21 (i.e. the fillet 40 does not extend around to the suction side surface 36 of the airfoil 21, such that the suction side 36 is free of any fillet 40).

The fillet 40 is disposed at the leading edge 30 of the airfoil 21, and extends at least partially along the leading edge a span-wise distance away from the root 26 of the airfoil 21. In one particular embodiment, the fillet 40 extends radially away from the root 26 a span-wise distance H of less than 10% of the total span S of the airfoil 21. More particularly still, the span-wise distance H of the fillet 40 may be from 2 to 10% of the total span S of the airfoil 21, beginning at the root 26 thereof.

As best seen in FIGS. 2 and 4, the fillet 40 of the depicted embodiment axially extends upstream and/or downstream relative to the leading edge 30 of the airfoil 21. It is understood that the leading edge 30 as used in this context (i.e. as a stream-wise reference point for the fillet 40) corresponds to the leading edge 30 of a majority of the airfoil 21 of the vane 20 outside the region of the vane having the fillet 40. More specifically, if a radially extending axis 50 is axially (i.e. stream-wise) aligned with the leading edge 30 of a major portion of the airfoil outside the fillet region, this defines the location of the “leading edge 30” as used as a reference point for the axially extending extent of the fillet 40.

The fillet 40 may accordingly extend upstream and/or downstream relative to the leading edge 30 of the airfoil 21, on the pressure side 34 of the vane 20.

As seen in the embodiment of FIGS. 2 and 4, the fillet 40 axially extends upstream a distance A from the leading edge 30 (and the axis 50), and axially extends downstream a distance B from the leading edge 30 (and the axis 50).

In one embodiment, the fillet 40 is disposed only on an upstream half of the pressure side surface 34, and therefore the distance B may be less than 50% of the total chord length C. In a more particular embodiment, the distance B may be from 10% to 50% of the total chord length C.

In one embodiment, the fillet 40 is disposed only within a radially innermost portion of the airfoil 21, proximate the root 26, and therefore the distance H may be less than 10% of the total span S. In a more particular embodiment, the distance H may be from 2% to 10% of the total span S.

In one embodiment, the fillet 40 extends upstream of the leading edge 30 (and therefore the axis 50) a distance A that is less than 20% of the total chord length C. In a more particular embodiment, the distance A may be from 5% to 20% of the total chord length C.

Accordingly, the fillet 40 in effect extends the leading edge of the airfoil 21, at its root 26, upstream (e.g. by 5-20% of the airfoil chord at the hub) relative to a remainder of the airfoil outside the fillet region. This has the effect of increasing the stagger angle at the root 26 of the airfoil 21 of the vane 20. The stagger angle θ is the angle defined between a horizontal axis X and a camber line CL of the airfoil, at any given span-wise location on the leading edge 30.

As best seen in FIG. 4, the stagger angle θ_(F) at the fillet 40 is greater than a stagger angle θ_(A) of a majority of the airfoil 21 at the leading edge 30 outside of the fillet region. More specifically, the stagger angle θ_(F) of the fillet 40 may be between 5 and 10 degrees greater than the stagger angle θ_(A) of the airfoil 21 outside the fillet region. By increasing the stagger angle θ_(F) at the leading edge of the root 26 of the vane airfoil 21, due to the presence of the fillet 40, the incidence angle of the airflow F is reduced (relative to that over the remainder of the airfoil, outside the filleted region).

The fillet 40 may accordingly help improve aerodynamic performance and flow quality downstream of the stator 17,17A, and may increase the stall range of the compressor 13. A reduction in stator flow separation may also result, which can lead to performance improvements for the downstream rotor(s).

Referring still to FIG. 4, the fillet 40 may have a fillet radius R that is from 5% and 20% of the total span S of the airfoil 21.

As also best seen in FIG. 4, in order to create the fillet 40, the suction side surface 36 of the airfoil 21 is extended upstream (from the leading edge 30), to thereby form the higher stagger angle θ_(F). The suction side surface 36 is extended smoothly upstream along the suction surface by the distance A, which thereby increases the higher stagger angle θ_(F) relative to the stagger angle θ_(A) of the airfoil 21 outside the fillet region.

As can be seen in FIGS. 5A and 5B, the fillet 40 on the pressure side of the airfoil 21 may be shaped with a blended curve (i.e. a so-called “variable” curve fillet), so as to smoothly (in an aerodynamic sense) blend the leading edge extension formed by the fillet 40 smoothly back to/with the pressure side surface 34 and the leading edge 30 of the airfoil 21 and the inner hub of the stator. This variable fillet shape on the pressure side may help to speed up the pressure side airflow, and reduce secondary flow losses. For example, this may reduce transverse flow, which results in less surface boundary layer radial flow migration. This “variable” or blended fillet 40 accordingly helps to speed up the pressure side flow and reduce secondary flow. The vanes 20 having this fillet 40 may also cause less transverse flow, which can result in less surface boundary layer radial flow migration. Overall stage performance of the compressor 13 may as a result be improved (lower stator losses and better aerodynamic performance).

The above description is meant to be exemplary only, and one skilled in the art will recognize that changes may be made to the embodiments described without departing from the scope of the invention disclosed. Still other modifications which fall within the scope of the present invention will be apparent to those skilled in the art, in light of a review of this disclosure, and such modifications are intended to fall within the appended claims. 

1. A compressor for a gas turbine engine, the compressor comprising: a rotor and a stator located downstream of the rotor, the stator comprising a plurality of vanes, each of the vanes having an airfoil extending generally radially from a root proximate an inner hub of the stator to a radially outer tip, a span of the airfoil defined between the root and the tip, the airfoil having a leading edge, a trailing edge, and a chord extending between the leading edge and the trailing edge to define a chord length, the airfoil having a pressure side surface and a suction side surface each extending on opposite sides of the airfoil between the leading edge and the trailing edge, and a fillet disposed at the root of the airfoil, the fillet extending upstream from the leading edge of a remainder of the airfoil, the fillet extending away from the pressure side surface a greater distance than the fillet extends away from the suction side surface.
 2. The compressor as defined in claim 1, wherein the fillet is disposed only on the pressure side surface of the airfoil, the suction side surface being free of the fillet.
 3. The compressor as defined in claim 1, wherein the fillet extends downstream from the leading edge.
 4. The compressor as defined in claim 3, wherein the fillet extends downstream from the leading edge a chord-wise distance of less than 50% of the chord length.
 5. The compressor as defined in claim 4, wherein the chord-wise distance is from 10% to 50% of the chord length.
 6. The compressor as defined in claim 1, wherein the fillet extends upstream from the leading edge a chord-wise distance of less than 20% of the cord length.
 7. The compressor as defined in claim 6, wherein the chord-wise distance is from 5% to 20% of the chord length.
 8. The compressor as defined in claim 1, wherein the fillet extends radially from the root along the leading edge.
 9. The compressor as defined in claim 8, wherein the fillet extends radially away from the root a span-wise distance of less than 10% of the span of the airfoil.
 10. The compressor as defined in claim 9, wherein the span-wise distance is from 2% to 10% of the span of the airfoil.
 11. The compressor as defined in claim 1, wherein a first stagger angle is defined at the fillet, and a second stagger angle is defined at the leading edge of the airfoil at a point thereon outside of the fillet, the first stagger angle being greater than the second stagger angle.
 12. The compressor as defined in claim 11, wherein the first stagger angle is from 5 to 10 degrees greater than the second stagger angle.
 13. The compressor as defined in claim 1, wherein the fillet has a fillet radius that is from 5% to 15% of the span of the airfoil.
 14. The compressor as defined in claim 1, wherein the fillet blends smoothly with the pressure side surface and the leading edge of the airfoil, and the inner hub of the stator.
 15. A turbofan engine comprising a fan and a casing defining a bypass duct surrounding an engine core defining an annular gas passage, a fan stator disposed within the engine core downstream of the fan, the fan stator including a plurality of vanes circumferentially spaced-apart around a circumference of fan stator within the annular gas passage, each of the vanes having an airfoil extending between a root and a tip spaced apart by a span of the airfoil, the airfoil having a leading edge and a trailing edge spaced apart along a chord line by a chord length of the airfoil, a pressure side surface and a suction side surface respectively extending on opposite sides of the airfoil between the leading edge and the trailing edge, and a leading edge fillet disposed at the root of the airfoil on the pressure side surface.
 16. The turbofan engine of claim 15, wherein the fillet extends upstream from the leading edge a distance of from 5-20% of the chord length.
 17. The turbofan engine of claim 15, wherein the fillet extends downstream from the leading edge a distance of from 10-50% of the chord length.
 18. The turbofan engine of claim 15, wherein the fillet extends radially away from the root a distance of from 2-10% of the span.
 19. The turbofan engine of claim 15, wherein a first stagger angle is defined at the fillet and a second stagger angle is defined at the leading edge of the airfoil at a point thereon radially above the fillet, the first stagger angle being from 5 to 10 degrees greater than the second stagger angle.
 20. A gas turbine engine comprising: a compressor with a rotor and a stator located downstream of the rotor, the stator having a plurality of vanes each having an airfoil extending span-wise between a root proximate an inner hub of the stator and a tip, the airfoil extending chord-wise between a leading edge and a trailing edge, a fillet disposed at the leading edge of the root of the airfoil and extending between a pressure side surface of the airfoil and the radially inner hub. 